1. Field of the Invention
This invention relates to deployable solar cell arrays, and more particularly to a semi-rigid solar cell panel whose width when deployed is in excess of the width of the spacecraft which stores the panels during launch.
2. Description of the Prior Art
There have been a variety of mechanical configurations proposed for deployable solar arrays. The various configurations fall into two general classes; rigid substrate and flexible substrate. The substrate is the structural element to which the rather fragile solar cells are affixed. In the case of the rigid substrate type of solar arrays, the structural element is usually formed of honeycombed material, and it is necessarily designed to be sufficiently rigid so that it is self-supporting between restraints during the launch phase. The conventional approach employs honeycomb material on the order of one inch in thickness for large arrays. No supplemental structure except for the deployment mechanism is required after deployment in orbit.
In order that an array with a large surface area may be stored within the confines of the launch vehicle shroud, multiple hinging of separate substrate panels is frequently employed. In such a concept, the several substrate panels with hinges between each panel are folded in accordion pleated fashion with the panels stacked one on top of the other with the stack extending radially outward of the spacecraft axis corresponding to the launch vehicle. Obviously, the height or radial extent of the stack of these panels becomes an important limitation because of the thickness of the substrate. Either the spacecraft volume is reduced to permit the necessary stack height, or the maximum power available is limited by the spacecraft volume requirements. Further, the panel width must decrease as the panels are stacked radially outward of the spacecraft axis, since the launch vehicle shroud being cylindrical in form must encompass the spacecraft and the separate rigid substrate panels of the solar cell array. Thus, the outermost panels must have a different width than the innermost if the deployment area is to be maximized, while shroud volume consumption is minimized. This not only introduces great complexity in the deployment mechanism, but requires a coordinated deployment sequence involving in most cases a scissors linkage attached to the edges of the substrates. Further, latching devices are required at the hinges between substrate panels and/or the linkage mechanism to rigidize the array, particularly subsequent to deployment.
The nature of any stacked panel solar cell array limits the maximum solar cell area presented for power generation prior to deployment, to one-half of the total area when two panels are used, one-third when three are used, etc.
Further, attempts have been made to place various hinged panel sections about the various sides or faces of the spacecraft body while maintaining the panel sections within the cylindrical volume limits of the launch vehicle shroud and while, at the same time, exposing the complete surface area of the panels prior to deployment. This keeps the deployment mechanism simplified, but these requirements require at least two hinged sub-panels affixed to primary panels if large solar cell surface areas are to be achieved.
Regardless of the nature of mounting the solar cell panel sections to the periphery of the spacecraft body, the rigid substrates, being relatively thick, constrains the maximum area by the fact that each panel element of the array lies on a chord of the circular shroud envelope. In addition, the chordal nature of the panel elements prohibits maximum spacecraft volume from being achieved. The deployed solar cells of the rigid class thus described suffer the inherent inefficiencies because of the discontinuous nature of the substrate. Each of the several separate panels or sections must have a portion of the substrate clear of solar cells around its periphery, about the points of hinge mechanization, and in the regions of the launch support interfaces. This, of course, requires more substrate area and weight than is necessary for the power function alone. In addition every discontinuity in the substrate surface imposes design limitations upon the electrical configuration of the solar cell layout, thus leading to additional inefficiencies. Further, at every hinge line, it is necessary to strap wire across the joint to provide electrical continuity and the system suffers the attendant reliability reduction. The nature of the rigid solar cell panels of the past design require that the necessary power area be achieved in finite increments with each increment the area of one sub-panel or section. For example; if 30 square feet of array is required and a design using two sub-panels is optimized, then each panel must necessarily be 15 square feet in area. Any growth in the requirements may only then be accomplished by adding a third panel having a similar 15 square foot surface area, not likely to be an optimum, or by major redesign of the sub-panels and launch restraint. Similar difficulties exist but perhaps to a more limited degree with the second of the above described type of fixed rigid solar cell panel array.
With respect to the second class of deployable solar cell arrays using the principle of a flexible substrate, the flexible substrate solar cell arrays are characterized by a "paper thin" substrate to which the solar cells are mechanically affixed. The negligible stiffness of this assembly requires special packaging to survive launch vibrations in a mechanized structure, to which the solar array is attached, for deployment and to provide sufficient structural stiffness in the deployed state. Solar cell arrays of this class are preferred over rigid substrate arrays, at higher power levels where they become weight competitive or superior and/or where the absolute power requirement (deployed area) exceeds that which can be realistically achieved with the simpler rigid substrate array. The substantial mechanical complexity of these arrays and the more severe transient temperatures (due to low thermal mass) suffered in orbit is generally believed to make them less reliable and more expensive than rigid arrays.
The present invention is directed to the solving of the problems inherent to conventional rigid substrate solar cell and flexible substrate solar cell arrays.